It is well-known that nonuniform tip clearance in an axial compressor induces pressure and velocity perturbations along the circumferential direction. This study develops a numerical modeling to predict perturbed flows in an axial compressor with a nonuniform tip clearance and presents a mechanism of the flow redistribution in the axial compressor at design and off-design conditions. The modeling results are compared with CFD results (2006, “Prediction of the Fluid Induced Instability Force of an Axial Compressor,” ASME FEDSM 2006, Miami, FL) not only to validate the present modeling, but also to investigate more detailed flow fields. In an axial compressor, nonuniform tip clearance varies local flow passage area and resultant axial velocity along the circumferential direction. There are small axial velocity differences between maximum and minimum clearances near the design condition, while large pressure differences are investigated according to local locations. However, contribution of the main flow region overrides the tip clearance effect as the flow coefficient deviates from the design condition. Moreover, the flow field redistribution becomes noticeably strong when the off-design effects are incorporated. In case of high flow coefficients, the low relative flow angle near the minimum clearance regions results in a large negative incidence angle and forms a large flow recirculation region and a corresponding large amount of loss occurs near the blade pressure surface. It further promotes strong flow field perturbations at the off-design conditions. The integration of these pressure and blade loading perturbations with a control volume analysis leads to the well-known Alford’s force. Alford’s force is always negative near the design condition; however, it reverses its sign to positive at the high flow coefficients. At the high flow coefficients, tip leakage flow effects lessen, while increased off-design effects amplify blade loading perturbations and a steep increase in Alford’s force. This study enables that nonuniform flow field, and the resultant Alford’s force, which may result in an unstable rotor-dynamic behavior, can be easily evaluated and assessed during the compressor, fan, or blower design process.

Several studies have previously been conducted to improve the cooling performance of film cooling. However, most of the research has conducted experiments with film cooling holes on flat plates, and thus, the results of these studies do not encompass the influence of the complex mainstream behavior within the turbine passage on film cooling. In this study, three different film cooling hole configurations were installed on the endwall of a turbine linear cascade to measure adiabatic film cooling effectiveness and evaluate cooling performance. The film cooling holes compared in the experiment for film cooling effectiveness were a 7-7-7 fan-shaped hole (Baseline), a Baseline with a double-step structure at the hole exit (Staircase), and a Baseline with an additional expanded passage at the hole leading edge (Compound Expansion). A total of nine holes were manufactured on the turbine endwall to assess film cooling performance, as various factors, such as mainstream acceleration, secondary flow within the turbine passage, and so on, can influence film cooling. Adiabatic film cooling effectiveness was measured using the pressure-sensitive paint (PSP) technique. Mass flow ratios ranging from 0.25% to 1.25% of the mass flow rate of a single turbine passage were supplied to the plenum chamber within the test rig. As a result, all experimental results confirmed the impact of secondary flow within the turbine passage on film cooling. In the case of the Staircase, it exhibits an overall cooling trend similar to the Baseline. It shows small cooling performance degradation compared with Baseline due to lift-off, and its double-step structure laterally expanding results in better cooling performance at high mass flow ratio (MFR) conditions. For the Compound Expansion, at low MFR, the momentum of the coolant is lower compared with other configurations, leading to lower cooling performance due to the influence of secondary flow. However, at high MFR, the Compound Expansion provides wider protection compared with other hole geometries and shows high cooling performance.

The objective of this study is to implement a more realistic integrated analysis model for micro gas turbines by incorporating secondary flow and combustion efficiency into the existing model, which includes main engine components such as the compressor and turbine, and to validate this model by comparing it with test results. The study was based on the JetCat P300-RX, which has a maximum thrust level of 300 N. Simulations were performed using ANSYS CFX, employing the κ-ω SST turbulence model and a mixing plane interface between individual components. The eddy dissipation model (EDM), with a combustion efficiency of 90%, was used as the combustion model. A user subroutine was also applied for the power matching of the compressor and turbine to calculate the fuel flow rate in each iteration. For secondary flow, it was assumed that 3% of the total air flow rate would flow through the secondary path and be applied to the compressor and turbine. Simulations were conducted over a range of 30,000 to 104,000 RPM, with ground conditions evaluated, including altitude-simulated conditions. To validate the analysis model, engine performance metrics such as pressure ratio, air flow rate, fuel flow rate, and exhaust gas temperature (EGT) were compared with test results. The results demonstrated that errors were less than 5% for most engine performance metrics, except for EGT and fuel flow. The discrepancy in EGT was attributed to differences in the sensing methods, while the variation in fuel flow was found to be due to the lubrication system and losses due to the secondary air flow. Consequently, this study confirmed that the integrated simulation model accurately predicts engine performance. The results indicate that the integrated simulation model provides a more realistic prediction of overall engine performance compared to previous studies. Therefore, it can evaluate detailed thermo-fluid properties without the need for component performance maps, enhancing performance evaluation and analysis.

This study presents the arrangement design of the cooling hole and the application of a fan-shaped hole to improve the cooling performance of the high-pressure turbine. To this end, the first stage nozzle vane of the energy efficient engine (E3) was selected as the base model, and for efficient design, the source term method, called the injection region, was applied for producing the effect of the cooling flow when the RANS analysis was performed. At this time, because the cooling flow rate also changed when the location of the cooling hole changed, a neural network model was constructed to predict the cooling flow rate for the location of the cooling hole. Design optimization was performed to improve the film cooling effectiveness and the temperature uniformity on the vane surface using the streamwise location of the cooling hole as a design variable, and then the cooling performance was investigated by applying several fan-shaped holes instead of cylinder holes on the pressure side. As a result, the final design was obtained, which improved the film cooling effectiveness by 4.1%p and temperature uniformity by 0.7% compared with the base model.